As a communication satellite moves around the earth, it passes through two phases in regard to charging and discharging of its on-board batteries. One phase is an eclipse phase, the other is a non-eclipse phase. In the eclipse phase, a solar array is not able to supply power to the spacecraft system and batteries are needed to supply the necessary power to the spacecraft. It is a challenge to keep the spacecraft sufficiently supplied with power by the batteries during the eclipse phase, as the output voltage of the batteries tends to decrease as the battery power or charge goes down.
It is known in the art that batteries used in spacecraft, specifically batteries used in a satellite, need to be light in weight and used most efficiently due to the launching cost and/or payload considerations. Different types of batteries have been developed to meet these needs. Batteries, such as nickel-cadmium, nickel-hydrogen or lithium-ion batteries, have been used. A typical satellite power system would include a set of batteries, a charge converter, and solar arrays all coupled to a common bus, with power being supplied alternatively by either the batteries or the solar arrays.
Spacecraft power system architectures most commonly comprise two types of systems, one using a direct energy transfer, fully regulated bus, and the other using a semi-regulated, direct energy transfer bus.
The direct energy transfer, fully regulated bus has the drawback of requiring a discharge regulator which is only utilized during a fraction of the spacecraft's operating life. Moreover, discharge regulators are expensive, heavy (which drives up launching cost), and they may include components interconnected in such a way that if some component(s) fail, the entire regulator will fail. Furthermore, discharge regulators consume some amount of spacecraft power, even during periods when they are not converting battery energy. In addition, the discharge converter must be operated at less than 100% efficiency, thereby effectively wasting battery energy.
The semi-regulated, direct energy transfer bus operates without a discharge regulator. While eliminating the drawbacks of the discharge regulator as listed above, the semi-regulated direct energy transfer bus has a drawback related to a non-optimum use of the solar array energy. The non-optimum use is caused by an effort to eliminate or avoid an undesired condition known as "solar array latch-up", which happens when the solar array output voltage becomes clamped to the battery. This typically occurs when the bus is changing from being powered by the battery to being powered by the solar array.
As the satellite emerges from eclipse, the solar array as well as battery begin to share the satellite load (both provide current at a battery voltage). This in effect reduces the battery current which in turn allows the battery voltage to rise. By example, with a negative impedance constant power satellite load, the rising battery voltage further reduces current demand. If the solar array is sufficiently large, its current will completely supplant the battery current, i.e., the battery ceases to discharge. Excess solar array current may now be used to charge the battery and the bus is deemed to be unlatched from the battery.
If insufficient array current is available, the battery will never fully return to open circuit. The solar array current becomes clamped by the battery voltage which continues to decrease as the battery is discharged, and a latch-up condition occurs. Without intervention, latch-up may lead to the loss of all satellite power.
In the prior art, the latch-up condition is avoided with an unregulated bus by over sizing the solar array such that the product of the array current and the open circuit satellite voltage at its maximum depth of discharge (end of eclipse) is greater than the end of eclipse bus power requirement. This approach, however, requires sizing the solar array for the end of eclipse power which is dictated by the battery voltage. This voltage is typically 65% of an equilibrium bus voltage. Thus, the array may be as much as 35% oversized with respect to its other operational requirements.
Typically, in a geosynchronous satellite power system, there are at least two sources of power. Referring to FIG. 3, a solar array 3 converts radiant energy into electrical energy for use by the satellite, and a battery 2 powers the satellite when the solar array 3 does not generate enough power to satisfy the requirements of the satellite. The prior art conventionally uses just enough battery cells such that when the battery 2 are fully charged, it does not exceed the maximum voltage allowed by a power bus 1. (see FIG. 3) Since the battery 2 is directly connected to the bus 1, and an unduly high voltage from the battery 2 is undesirable, the choice of the number of batteries is limited by a maximum voltage allowed by the power bus 1.
For example, if nickel-hydrogen(NiH) cell batteries are used in a space craft power supply system, the maximum number of Ni-H.sub.2 cells used may be about 60 cells for a nominal 100 volt voltage level. Nickel-hydrogen batteries can be operated at a maximum of 1.65 volt per cell(at the end of battery charge), and a minimum of 1 volt per cell (at the end of a worst case discharge). Thus, the maximum voltage would be 60.times.1.65 volt/cell=99 volts, and the minimum voltage would be 60 volts. In a typical unregulated high voltage bus system, this will result in a bus variation(or range) of 60-99 volts, or 39 volts. This wide operating range can result in solar array latch-up at the end of the eclipse period, and also tends to reduce the efficiency of the power converters during periods of battery operation.